1. Field of the Invention
The present invention relates generally to spacecraft and, more particularly, to spacecraft attitude determination and control.
2. Description of the Related Art
Points on the earth do not move relative to a geostationary spacecraft. Accordingly, the spacecraft can be maintained in an appropriate service attitude and the positions of earth points, relative to that attitude, will remain constant over each solar day. An exemplary geostationary spacecraft is a communication spacecraft that provides a payload beam which serves a communication service area on the earth and facilitates communication between points in the communication service area and the spacecraft.
In particular, the payload beam is configured to define a payload footprint on the earth that is preferably identical to the communication service area. Because earth points remain fixed relative to the service attitude of a geostationary spacecraft, the payload footprint remains substantially fixed over each solar day. This characteristic of geostationary spacecraft facilitates minimization of service error which is any difference between the payload footprint and the service area.
In contrast to a geostationary spacecraft, FIG. 1 illustrates a communication spacecraft 20 that orbits the earth 22 in an orbit whose orbital plane 24 is inclined by an inclination angle 26 from the earth""s equatorial plane 28. The spacecraft carries an antenna system 30 and solar wings 32 and is shown in its service attitude at positions 34A, 34B and 34C which correspond to times TO, TO+6 hours and TO+12 hours. The spacecraft may use signals from a beacon station on the earth 22 as an attitude reference.
Because the spacecraft 20 of FIG. 1 is in an inclined orbit, earth points will move relative to the spacecraft""s service attitude. In particular, they move along a figure-eight path such as the path 40 of FIG. 2A which indicates movement direction by a path arrowhead. Earth points initially drift southward and eastward during the first quarter of the orbit from its ascending node (spacecraft at position 34B in FIG. 1). For example, an inclination (26 in FIG. 1) of 5.4 degrees will cause an earth point (e.g., a beacon station) at approximately 25xc2x0 north latitude to trace a path having a north-south angular extent 42 on the order of 0.72 degrees and an east-west angular extent 43 on the order of 0.14 degrees.
More generally, the path 40 will define distorted figure-eight patterns that are tilted from a north-south axis 45 as shown in FIGS. 2B and 2C. The path 40 and the area within the path can be considered to define a beacon-station window 44 (i.e., a window, as observed from the spacecraft, that always contains the beacon station).
Various optimal steering laws have been utilized to realize different spacecraft pointing objectives (e.g., see U.S. Pat. Nos. 5,184,790, 5,738,309 and 6,135,389). A spacecraft""s service attitude is determined by its respective steering law and the motion of an earth point, relative to the spacecraft, is a function of the steering law, orbital eccentricity, spacecraft mean longitude error, orbit inclination and longitude/latitude of the earth point relative to the spacecraft""s location.
The service attitude of a communication spacecraft is typically maintained with an attitude-control system that receives attitude input signals from various attitude sensors. An exemplary set of attitude sensors comprises a sun sensor and a beacon-receiving antenna that receives a beacon signal from a beacon station on the earth. The beacon-receiving antenna is typically realized with several similar antenna beams that are arranged in a pattern such as the three-beam pattern 50 of FIG. 3.
In this pattern, the beam widths of the three beams are represented by similar circles 51 which define beam points that have a common power level and are arranged to intersect at a beacon-receiving boresight 52. Other exemplary beacon-receiving beam patterns in FIG. 3 are the four-beam pattern 54 and the four-beam pattern 56 which also define beacon-receiving boresights 52. Because each beam""s power slope increases off-peak, the beam patterns of FIG. 3 form sensitive beacon-receiving antennas.
The field-of-view of a beacon-receiving antenna is defined as the area over which it provides a useful attitude signal and is substantially determined by signal-to-noise considerations. The three-beam pattern 50 forms a substantially-triangular field-of-view 53 and the four-beam pattern 54 forms a substantially square field-of-view 55. The four-beam pattern 56 includes a first pair of beams 57 that are alternated with a second pair of antenna beams 58. Because each of the second pair has a beam width that is substantially broader than that of each of the first pair, the four-beam pattern 56 forms an elongate field-of-view 59. Typically, the fields-of-view of beacon-receiving antennas (e.g., 53, 55 and 59 in FIG. 3) have been enlarged to a size that insures they will contain the beacon-station window (e.g., 44 in FIGS. 2B and 2C) throughout a communication spacecraft""s orbit.
FIG. 4 is a view of elements within the curved line 4 of FIG. 1. FIG. 4 illustrates that the antenna system 30 generates a payload beam 60 that has a payload vector such as the payload-beam boresight 62 fixed in the payload beam. The payload beam illuminates a payload footprint 64 on the earth 22 that is preferably identical to a communication service area. It has been found that the service error (between the footprint and the communication service area) is reduced if the payload vector is directed over each solar day at a subterranean target 66 as taught, for example, in U.S. Pat. No. 6,135,389.
The spacecraft 20 has two attitude sensors in the form of a sun sensor 68 and a beacon-receiving antenna that is realized with the antenna system 30. The sun sensor 68 is preferably one having a wide field-of-view (e.g., 120xc2x0) that can provide an attitude signal for a significant portion (e.g., 4-6 hours) of each solar day. The spacecraft""s attitude control system preferably includes a gyroscope system that estimates attitude about the yaw axis for the remaining portions of the day. With attitude input signals from the sun sensor and the gyroscope system, the spacecraft""s attitude control system is able to control the spacecraft""s attitude about its yaw axis which is generally coaxial with the payload-beam boresight 62.
The beacon-receiving antenna (part of the system 30) has a beacon-receiving boresight 70 which generally differs from a beacon line-of-sight 72 from the spacecraft 20 to a beacon station 74 which radiates a beacon signal. The beacon-receiving antenna provides a difference signal which corresponds to the difference angle 75 between the beacon-receiving boresight 70 and the beacon line-of-sight 72. The difference signal is a useful attitude signal over the beacon-receiving antenna""s field-of-view (e.g., 59 in FIG. 3). With the difference signal and yaw information from the sun sensor and gyroscope system, the spacecraft""s attitude-control system is programmed to direct the payload-beam boresight 62 at the target 66 over each solar day.
The payload beam is typically formed with a plurality of spot beams (e.g., on the order of 200). Reduction of service error has typically been realized by uploading beam coefficients and beam weights throughout the solar day that appropriately steer and reshape the spot beams. The specifications of many modern communication systems, however, are quite demanding and these systems have generally observed that the service error remains excessive and further reduction of service error would be useful.
The present invention is directed to methods and structures that enhance service attitude accuracy of inclined-orbit spacecraft and, thereby, facilitate reduction of service error between a communication service area and a spacecraft""s payload beam.
These goals are realized by configuring a beacon-receiving antenna to have a beacon-receiving field-of-view that substantially matches a beacon-station window (a window, as observed from the spacecraft, that always contains a beacon station). Preferably, the beacon-receiving field-of-view is elongated and tilted to enhance its match with the beacon-station window in both size and orientation.
The goals are also realized by configuring the beacon-receiving antenna to have a beacon-receiving field-of-view that is substantially smaller than the beacon-station window and successively steering a beacon-receiving boresight to successive beacon-receiving attitudes that maintain the beacon station within the beacon-receiving field-of-view over each solar day. In an embodiment of the invention, successive positions of the beacon-receiving field-of-view are arranged in a tiled arrangement.
Spacecraft structures are also provided to practice the methods of the invention.
The novel features of the invention are set forth with particularity in the appended claims. The invention will be best understood from the following description when read in conjunction with the accompanying drawings.